text:
Can one trim same AOA at a higher airspeed to climb?
There has been some discussion in recent questions regarding climbing technique and optimization of the Lift vector (and the resultant vertical Lift vector).
Some writers claim the lift vector must be smaller in a climb.
This does not seem to make sense, as the wing is more efficient than the propeller at generating force, by a factor of 4, for an average GA aircraft.
Although there is more drag at higher speeds, is it not better to climb at a higher speed with a larger Lift vector than it is to pitch up and use the engine thrust to climb?
Climb and distance Vy vs Vx, and glide distance Vbg vs Vmin sink seem to support this concept.
Can I maintain optimal Lift/Drag AOA at a higher airspeed by reducing elevator trim?

"Although there is more drag at higher speeds, is it not better to climb at a higher speed with a larger Lift vector"-- why assume that more airspeed corresponds to a larger Lift vector?  This is only true if the angle-of-attack is constrained to be constant.  Such a constraint is not consistent with the basic flight dynamics of free flight.  I think the question can be improved.  It's hard to be sure exactly what you are asking, other than questions that have already been asked like "Is Lift larger than Weight in a climb?"

We need a question about "what does elevator position most directly control: angle-of-attack, or pitch attitude?"  There may already be one.

Anyway the answer to the last sentence is "no, starting at a given weight and a given airspeed, and a given thrust setting, if we change the elevator position to improve the L/D ratio, the airspeed will change as well, as will the climb angle."  Does that help?  Waiting till question is stabilized to post as an actual answer.

"This is only true if AOA is constrained as constant" is the question!

Sorry to aggravate you; just you know it wasn't my dv.  I've sworn off dv'ing on ASE.

@quiet flyer appreciate that. I'm from R&D background, so I'm used to "dvs". For the 737, was thinking if replacing the engines with fuel pods (like the P-51) to make it staticly stable on takeoff (jets in back). The fuel would drain in cruise.  Not sure if that idea will make it either, but it's fun to try.

So you want the CG to move aft as fuel is consumed?  You can only let it shift to the aft CG limit which must remain forward of the neutral point at all times, so that tail downforce is required to oppose a ND pitching moment that's always there, or else you become unstable in pitch and once the CG is aft of the NP you get pitch instability and trim reversal (to slow down requires more ND trim, not more NU trim).  You can't just move wing mounted engines to the tail and offset it temporarily with fuel.

I never DV posts either.  This isn't PSE, which is a political cesspool dominated by a faction that bullies dissenters off the channel.   Unless a post is obviously a troll or someone looking for others to do their homework for them, any question is good with me and if I don't like a question I just ignore it.

Please, what is "DV"?

DV or dv = "downvote"

@John K correct, it would take a bit more than that.  It did occur to me the slightly forward fuel tanks on the P-51 made it bit more slow and docile, until they were dropped.  Actually, my horrible question does have an answer that applies to gliders too: yes, if the aircraft is heavier, it will have a higher speed with the same AOA.

Mustangs took off on long range missions with the aft tank full in addition to drop tanks and the CG very close to the NP and were very unpleasant to fly, nearly neutral in pitch. The tank behind the seat was always burned off first to get rid of the marginal pitch stability and low stick forces before anything was taken from the DTs.  Dogfighting with fuel in the fuselage tank was very dangerous because the stick force per G was much too shallow and it was easy to overstress, or enter into an accelerated stall if below Va.

Aft tank was burned off first and what do you have?

Wings and drops.

Here is one of those writers who claim that lift is smaller in a climb than in straight flight.
With good reason.
The sum of all vertical forces will be zero in a stationary climb or the aircraft would accelerate. It is a fallacy to assume that lift is higher at higher speed – lift is just as large as is needed to equalize all other forces. Now it is important to note that the aerodynamic coordinate system (in which lift is defined as the aerodynamic force parallel to the Z axis) is tilted, pointing slightly upwards with its X axis parallel to the flight path. Drag is slightly pointing down and lift is slightly pointing back when seen from an earth-fixed observer.
To equalise those forces, we have only thrust (at least in the absence of wind – or this would be the special case of a glider climbing in a thermal). This thrust must now compensate both the drag force and the backward component of lift. It is, therefore, higher by that backward component of lift compared to straight and level flight. And it is also pointing slightly up due to the climb attitude of the aircraft. Therefore, this upward-pointing thrust component (which is larger than what is needed to compensate the downward pointing drag component) works against gravity and helps the lift in compensating weight. Lift can, therefore, be a bit smaller than in straight and level flight.

Forces and their angles acting on a climbing aircraft in side view. The light, translucent arrows are copies of the solid arrows shifted into a vector chain to show that they indeed equalize to zero net force.
The extreme case is a vertical climb: Now all the force counteracting weight is supplied by thrust and lift is zero, or the aircraft would accelerate sideways. Climb is simply a linear combination of level flight and a bit of vertical climb.
Now you could argue that the B-52 climbs with its nose pointing down, so thrust must also point down a bit. Maybe, but in level flight that thrust would point even more down and more lift would be required to compensate for that downward thrust component. No matter how you look at it, as soon as the flight path angle pitches up, so does the thrust, thereby reducing the required lift.
Now you argue that climbing at a faster speed is better. But that would create more drag, leaving less excess power for climbing. Also, when flying faster, the aircraft must reduce its angle of attack in order to produce the right amount of lift. Climbing is not about lift maximisation, but about excess energy maximisation.

What Peter is saying is climb is a function of thrust that is in excess of the minimum required for level flight and the maximum thrust surplus occurs at best rate, near maximum L/D.  Any faster is less efficient in terms of climbing from x to y altitude with the minimum additional energy.

Actually, sustained wind and updraft don't change the geometrical orientation or magnitude of the forces either as long as the airmass is locally uniform on a large enough scale for the glider to come to equilibrium.  The only affect the trajectory as seen from the ground.  Content in first set of parentheses could simply be omitted.
Really have to get that diagram added to my answer to related question.  Ditto to Robert's question about want to understand the vectors for Vbg.

No.  There is a direct correlation between airspeed and AoA.  In level, balanced unaccelerated flight you cannot change airspeed without affecting the angle of attack.
L/D max AoA will occur at a certain airspeed for a given weight.  It is not possibly to fly a different airspeed at the same AoA unless you are in accelerated flight.  (I.e. pulling Gs in a turn)

Can I maintain optimal Lift/Drag AOA at a higher airspeed by reducing
  elevator trim?

No, starting at a given weight and a given airspeed, and a given thrust setting, if we change the elevator position to change the angle-of-attack to change the lift coefficient and drag coefficient in a way that improves the L/D ratio, the airspeed will change as well, as will the climb angle or descent angle.  There is no other way to end up with closed vector polygon of Lift, Weight, Drag, and Thrust.  For a given thrust or power setting and a given weight, if you use the elevator to change angle-of-attack, you will end up with a change in airspeed.
For shallow climb or descent angles, in sustained linear flight, it's a good approximation to say that there's a one-to-one correlation between angle-of-attack and airspeed, or between angle-of-attack and L/D ratio (excluding the "back side" of the polar curve, or between airspeed and L/D ratio (excluding the "back side" of the polar curve).  Changing any one of these things will automatically change the others.
For very steep climb or descent angles it gets a little more complex because the amount weight supported by the wing is less than in level flight, so the airspeed associated with any given angle-of-attack and L/D ratio will be less than it was in horizontal flight, just as if we had changed the weight of the aircraft.  But even in a very steep climb or dive, for a given power or thrust setting, we can't choose between different L/D ratios (and different resulting climb or dive angles) for the same airspeed, unless we are changing the shape of the airfoil or aircraft by deploying flaps, spoilers, etc.
Nor can we have our pick of different airspeeds for the same angle-of-attack and thrust or power setting.


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Longitudinal stability of airships: How is the critical speed defined?
Airships are unstable in pitch above a certain speed, called the critical speed. What causes the instability, how does it manifest itself for the pilot, and how is the critical speed defined and determined?

I've got the feeling you are going to answer this one yourself

hmmmm, there's no Airship Flying Handbook to go along with Airplane, Glider, Helicopter, Rotorcraft (Gyro), and Instrument

Airships have a center of lift well above their center of mass, since engines and the gondola are mounted at the bottom. This gives them a strong static pitch stability: They will try to assume an attitude where the center of mass is right below the center of lift. Normally, the airship will be carefully trimmed to ensure that the hull is level at this attitude.

During flight both the mass and the lift will not stay constant: Sun might shine on the hull, heating and expanding the gas inside and increasing lift, the hull might be wetted by rain and become heavier, or fuel burn will lighten the ship. To stay at the desired altitude, the airship can compensate for any imbalance by creating dynamic lift. By pitching the hull up or down, up to 10% of the needed lift can be added or subtracted this way. Note that this will not only create aerodynamic lift, but it will also point engine thrust slightly up- or downwards, which will add to the lift change. If the ship is fast enough, a negative deflection of the horizontal tail surface or a shifting of masses backward will increase lift.

At low speed, however, both thrust and aerodynamic lift are much lower, but both mass and static lift are unchanged. Now the ship needs considerably more elevator deflection for the same pitch attitude change, but aerodynamic lift and engine thrust will be a lot less. The downforce at the horizontal tail must still be equal to that at high speed, since the static moment has not changed, and the lower dynamic pressure at low speed is compensated by deflecting the elevator by a  bigger angle.
Below the critical speed, the elevator downforce will become dominant, and pitching the ship up will decrease total lift. The critical speed is characterized as the speed at which attitude changes will not result in lift changes, and when crossing this speed, the elevator effect on lift reverses. For this reason some books use the term "reversal speed" for the critical speed. No airship captain wants to cruise at this speed, since he will be bereft of means to change lift quickly!
This effect can create a dangerous situation if a heavy ship comes in for landing. In cruise, the missing static lift could easily be compensated by raising the nose, but now the ship has to decelerate. It will start to pick up sink speed, and raising the elevator will make matters worse. The only way to avoid a crash in this situation is the application of reverse thrust combined with a positive elevator deflection which lifts up the tail. Proper timing is of the essence: The ship must touch down at the desired location just when the reverse thrust has stopped both the vertical and the horizontal speed. This requires a trained crew and a skilled commander. If the propellers cannot run in reverse, this situation must be avoided at all cost!
The critical speed is higher the blunter the airship is. The slender Zeppelins of the first World War had a fairly low critical speed (maybe 8 m/s), but the Cargolifter's critical speed was around 20 m/s. During the design phase the required power had to be increased by 40% to raise the cruise speed by a sufficient margin above the critical speed.
Both the vertical and the horizontal tail surfaces of airships are much smaller than what would be required for natural stability, and they typically were just big enough to keep the ship under control. Aerodynamic lift is mostly created at the forward part of the hull, and lifting the nose will increase the pitch-up moment. The same goes for the sideforce: In both axes the helmsman needs to constantly correct the control surface deflection to keep the ship at the desired attitude and heading. Below the critical speed, the aerodynamic forces and the pitch-up due to the low location of the propellers will not be sufficient to overcome the static stability, so pitch excursions will be limited. Above the critical speed the aerodynamic instability will become dominant and the helmsman needs to be alert and responsive.

While not exactly the same, there are two other areas that have a similar reversed control input effect. The flare at the bottom of a helicopter auto rotation is using the same force for lift to slow the craft in its decent. Where in powered decent you are normally lowering the collective. The other is prop-walk on a boat where at low velocity the propeller exerts side pressure on the rear of the vessel in the direction of prop rotation at speeds below those with rudder authority. It is more pronounced at higher throttle in reverse in that in forward but it is still there in both directions.

With the auto rotation case I do realize that you can over control and pop up and then crash but I mean that the control inputs are reversed from what you normally would be doing on a powered landing.

On page 43 of the Technical Manual of Airship Aerodynamics I found the following image:

Here $F_e$ is the resultant airforce on the hull, or the dynamical lift of the hull. As this is before the CoG, it is destabilizing.
This source mentions Munk's formula being:
$$ M_e = Volume \cdot 0.5 \rho V^2 (k_2-k_1) \sin{2\theta} $$
Where $k_2$ and $k_1$ indicate the effect of gas movement inside the airship.
I tried to trace the origin of this formula in Munk's paper, available here, but I can't immediately find a clear explanation.
The critical speed is reached when this resultant airforce equals the moment induced by the weight:
$$Wh\sin{(\alpha + \theta)}=Volume \cdot 0.5 \rho V^2 (k_2-k_1) \sin{2\theta}$$
In this sense, the critical speed is the maximum speed at which the gravity moment is able to compensate the moment induced by dynamical lift. As such, one should stay below the critical speed.

Thanks for the link to the Technical Manual! If you read on, on page 51 the topic is mentioned. What I called "critical speed" is called "reversing speed" in the manual, however.

Do you fully understand the reason mentioned? As I read it, it says that the moment induced by the elevator is quickly compensated by the static equilirbium moment, but the downward force of the elevator will push the airship down, is this correct?

Is this just the gas sloshing, like a long jar or tube with some water in it over a centered fulcrum?

@ROIMaison: Yes. Pretty simple, right? But not intuitive.

@rbp: Soft ballonets will indeed exacerbate a static imbalance, but the real reason for the instability is the low thrust at low speed.

I'm not sure if I should leave my answer here. It is an answer to 'what is the critical speed', but not in the way Peter meant. His own answer is a much better answer to his question.

@ROIMaison: Stackexchange thankfully allows to edit and expand answers. Yours has useful links and I encourage you to keep it here. If you are not satisfied, you may edit it any time.

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Significance of $\frac{C_L^{3/2}}{C_D}$
I am aware that this value of lift coeff. over drag coeff. occurs when the zero lift drag is 3X the drag due to lift, and that this will correspond to a specific airspeed. What is the significance of this value/ this airspeed?

Gliders fly at their minimum sink speed when $\frac{c_L^{3/2}}{c_D}$ reaches its maximum.
Propeller airplanes fly at their maximum endurance speed and their best climb speed when $\frac{c_L^{3/2}}{c_D}$ reaches its maximum.

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text:
Could simple fly-by-wire controls be installed in a prototype as quickly as a mechanical system with an autopilot?
It seems that fly-by-wire systems can take longer to get certified as the testing requirements are much more extensive. A prototype aircraft wouldn't have to pass these tests.
How much more time or complexity might it be to implement a fly-by-wire system for just one set each of ailerons, rudders, and elevators? An equivalent level of functionality in the mechanical system would require at least an autopilot.
Any links to regulations or example systems would be greatly appreciated.

This question is fairly broad as written but can easily be brought in a bit. What size prototype? An airliner? A GA Plane? Are you basing this of an existing system or asking about a clean sheet design for something that has not yet been tried? More time for who, a big manufacturer or someone building an experimental in the back hangar?

Early civilian FBW systems were prototyped by taking a conventional cable aircraft and feeding the control surface commands into the existing autopilot system and servos.
The earliest was with NASA Langley's 737 which had an aft "research" flight deck. Various instrumentation and control schemes were tested. The airplane's normal flight deck remained in the front with safety pilots.

(NASA Photo)
When Airbus developed the FBW system for the A320, they took a test A300 (F-BUAD), installed the sidestick on the left seat and rigged the FBW system to give commands via the autopilot system.

(Source, Article)
This was the last time this was done; no subsequent Airbus aircraft ever had cable prototypes and to my knowledge Boeing never did. FBW systems are mature enough that there does not need to be cable backup testing. If you need to prototype some new control law or scheme, you can simply modify the software in an existing FBW aircraft.
A fundamental issue is that the intent of testing is to build a new aircraft type. It takes quite a bit of engineering work to design a cable system that would only be used for early stage FBW testing, and would ultimately be wasted. In a modern FBW system, everything is completely tailored to electronic control, from the stick all the way to the hydraulic actuators.


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Why do we use dimensionless expressions in flight mechanics and aerodynamics?
Why do we use non-dimensional expressions in flight mechanics and aerodynamics? We could as well directly calculate forces and moments; would that not be more relevant to the specific problem?

Yes but in physics meaning what useful of it , you can use for example use force as dimensional formula but in flight mechanics we use dimensionless formula .

I have to disagree; the question is clear. We mention the subject of dimensionless coefficients in countless answers (see here for example), but do not yet have one that explains why this is done so. I vote for reopening the question.

Here is a dimensionful equation: $$L = \frac{\rho}{2}\cdot v^2\cdot \frac{2\cdot\pi\cdot b^2}{1+\sqrt{1+\left(\frac{b^2}{2\cdot S}\right)^2}}\cdot\alpha$$
Note that all ingredients are physical, measurable values. Now here is the same thing again, now in dimensionless form:
$$c_L = \frac{2\cdot\pi\cdot AR}{1+\sqrt{1+\frac{AR^2}{4}}}\cdot\alpha$$
The dimensionless version is independent of speed $v$ and air density $\rho$ and also needs only the aspect ratio $AR$ and the angle of attack $\alpha$ to give a result, while the dimensionful version needs wingspan $b$, wing area $S$ and angle of attack. However, while the dimensionful version yields an actual force, the dimensionless version yields a coefficient which needs to be multiplied with speed, density and wing area again in order to result in a force.
By working in the dimensionless domain, engineers can work on a generalized form of the problem. This was especially helpful in the time before computers: The complex calculation to arrive at a coefficient was needed only once, and computing the dimensionful value from there was trivial.
Also, when comparing between different airplanes, or between wind tunnel models and the real thing, with dimensionless coefficients the numbers will be very similar, which allows much better to error-check results or to make first assumptions on a new design.
However, there are two drawbacks:

You need to know the reference values! It is easy to screw up, e.g. when combining values from the wing and the tail which are referenced to their individual area. While lift can be added right away, the lift coefficients must first be converted to a common reference value.
Sometimes thinking in coefficients can be misleading: Just consider induced drag. If written in dimensionless coefficients like this
$$c_{Di} = \frac{c_L^2}{\pi\cdot AR}$$
you can be forgiven for thinking that induced drag goes down with increasing aspect ratio. But this is wrong! Now the same again with dimensions:
$$D_i = \frac{L^2}{\frac{\rho}{2}\cdot v^2\cdot\pi\cdot b^2}$$
This shows that induced drag at the same speed is inversely related to span loading, the amount of lift created per unit of wing span.

Some of the common values are also dimensionless by nature--for one obvious example, Reynolds number is computed as a ratio in which all the dimensions cancel out, leaving only a dimensionless number. As such, in quite a few cases, you simply can't avoid working with at least a few things that are dimensionless.

@JerryCoffin: But then the Reynolds number was defined in order to make different flow conditions comparable, just like the force and moment coefficients. The difference is that using viscosity, length and density explicitly would needlessly complicate the calculations, but it can be done.

I suppose you could do calculations using dimensional quantities, but it would do more than just complicate calculations--it would hide the applicability of the results (e.g., by knowing that they operate at close to the same Reynolds numbers, you know that calculations for the airfoil of a relatively slow-moving glider apply about equally to a jet engine's turbine blades).

Many thanks that was very helpful.

So that boils down to in engineering, we actually use both. Dimensionless expressions are handy e.g. when comparing two designs, while dimensionful expressions are used when designing. An example: when comparing the price of an A380 to a Cessna 172, it's good to know the A380 is 1032.4x more expensive. If i want to know, whether I can afford one, I'd want to know the price in $$$

As Stefan also wrote, in science your goal is to ignore e.g. the dimensions your experiment is build in to be able to compare it with other results (regarding similitude theory, which also uses dimensionless numbers).
With the gathered dimensionless data it is also much easier to develop theories independtly from any unit system. Furthermore, the variables of fluid dynamics are often connected with power laws. These are much easier to handle mathematically correct if you already got rid of the units.
Lastly, when you apply the theory to a specific engineering problem, you can reintroduce your dimensions etc. to get the result.

I would like to answer this question from a historical perspective.
Dimensionless numbers are used much more heavily in fluid dynamics than in other engineering disciplines. For structural mechanics, we happily throw around parameters like Young's Modulus, density, moments of inertia, and expect people to know what these numbers mean, even though they work out completely differently on different scales. We know approximately how things scale in simple objects, and whether we are calculating a landing gear or a wing, the beam bending problem is approximately the same.
Not so in fluid dynamics. Of course, Navier-Stokes, Bernoulli and friends had long ago figured the governing equations for fluid dynamics, but their applicability is much more limited than say, Hooke's law of elasticity. In structural mechanics, one can assume the wing to look approximately the same before and after bending, but in fluid dynamics the initial conditions and geometry tells us nothing about the final results, and a small perturbation on one side may completely change the results on the other side.
These days, we can use CFD (computational fluid dynamics, "computer simulation") tools to get from a set of initial conditions and (dimensionful) properties to a final result. In "the olden days", wind tunnels were used extensively. Since these things are expensive to build and maintain, and difficult to make on a large scale without computers, they were generally kept quite small. As a consequence, tools for scaling were very important - what do small-scale results tell us about the final design?
AS it turns out, dimensionless numbers are very useful for scaling. Since they do not depend on the unit system used, one can infer that they do not depend on the scale of the problem - the number should be the same whether we use a foot-long model or a meter-long one. Flow will become turbulent if $Re>2300$ (by appromation - you will find different values depending on application), and this tells us that, when we have a small model that we test near a Reynolds number of 1000 and all seems well, we may conclude that this also holds for the final airplane if the Reynolds number is around 1000 for that design.
You will note that dimensionless numbers are less interesting in the computerized age (why jump through all these hoops to get all the units out, if you can just plug it into the computer) - for disciplines where FEM etc. is ubiquitous, you will only see them used as a back-of-the-envelope calculation. In fluid dynamics, I should expect them to stick around a bit longer, until CFD becomes powerful enough to readily calculate large models within very short time spans.

As far as aerodynamics are concerned: with dimensionless coefficients (i.e. for lift and drag), statements independent of profile, body size and dynamic pressure can be made for every airfoil and its properties for different flow directions. That allows comparison of different profiles and airfoils - in flight mechanics or aerodynamics, oftentimes it is more interesting to compare airfoils and determine the best, instead of calculating exact forces).

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Is the climb rate different for short-haul flight and long-haul flight?
A short-haul flight does not spend much time in its cruise level (let's say less than two hours). Thus, it is important to rapidly reach the optimal flight level and I think this kind of aircraft should be optimised to climb. On the opposite, on long-haul flights there the climb and descent reprsent a tiny portion of the whole flight. The aircraft should then be optimised to consume as little fuel as possible during cruise, even if climb rate is less important. That's why I wonder if my thinking is right and if there are significant differences in climb rate between short and long-haul aircraft.
Related thinking: On average, does a short-haul flight descent more rapidly than a long-haul? I think not as this diving has no reason to be related to climb/cruise performance and all commercial aircraft are made to use the same airports with the same approach pattern; but that's only an assumption.
EDIT: OK, the quantity of fuel is quite important. I suppose an aircraft designed to be heavier (long-haul, carrying lots of fuel and payload) is equipped with more powerful engines and thus if it is lighter than designed for (lets say a B777 or a A330 used on short-haul flights), it climbs faster than the same aircraft in a long-haul flight. But the long-haul aircraft is designed to make long-haul flights. Maybe I should have asked if there is a significant difference in climb rate between short and long-haul aircraft at their maximum operating take-off weight.

The bigger difference is less fuel and consequently weight on the short haul flight that allows the plane to get higher quicker.

@ratchetfreak Yes, and I suppose the maximum takeoff wheight (including the fuel capacity) is a key figure to choose the total trust. A by-design lighter aircraft will have less powerful engines, conterbalancing the weight advantage for climbing.

@ManuH I think the interesting case is comparing the same aircraft on short and long hauls rather than comparing different aircraft where there are many other factors getting in the way of a clear answer. So maybe answerers could pick some specific aircraft and compare the different situations for them.

Different climb performance is indeed indirectly linked to the length of the trip, but this has to do with the number of engines. All airliner's engines are sized to support flying with an engine failure right after take-off. Twin-engined aircraft enjoy slightly less stringent rules, but will basically carry twice the thrust that they need to stay aloft, so more excess power is available than in four-engined aircraft, which have only 33% excess thrust.
Generally, the best performance is achieved when running the engines at their maximum continuous thrust setting and when climbing as fast as possible. That is the reason why an A-320 or a Boeing 737 will climb faster than an Airbus A-340 or a Boeing 747. On descent, the number of engines is less relevant, so the descent speeds of all types are similar.
The link between engine numbers and range was strongest in the early jet age; now it is less clear-cut with long-range airplanes like the Boeing 777-200 LR or short-range airplanes like the BAe 146.

No, not for the reasons you ask about.
Absent restrictions due to other traffic, terrain or weather, pilots will always choose to maintain the best rate of climb speed ($v_y$, a.k.a. "green dot speed") and run the engines at the recommended climb power, simply because that's when the plane flies most efficiently. The actual performance differs due to weight and weather, but that is not directly related to length of the flight leg.
Similarly the descent is again flown in a way that makes it most efficient (I believe that means close to maximum rate without speed brakes), which is again not directly related to length of the flight.
What short flights do is stop the climb at lower altitude (than permitted by load), because the time spent on higher level would be too short for the lower fuel consumption to offset the additional fuel needed to get there.

Short answer: yes.
Long answer: on a long haul aircraft, it carries a lot of fuel for the next 12 hours. That extra weight has quite an impact on climb performance. E.g. it takes more time for a fully loaded B777 to climb to 10000 feet, compared to a B777 configured for short haul. (Yes, many Asian airlines use B777 for short haul trips. The shortest one I can think of is CX flying Hong Kong to Taipei, a trip less than 2 hours)
Long haul flights also just cannot climb that high. The pilots can crank the engines max and it'll just go perhaps FL270 - FL300. Once after few hours, they burn up some fuel than climb, up to FL390 or higher. These are called step climbs : burn some gas, climb a bit, burn some more gas, climb again, and so on.

You mention different reasons than in the question, so I believe your short answer should really be "no".

Long haul flights also carry less passengers and cargo. I don't think weight is a big issue here as T/O weight for both flights will more or less be the same (One carries more pax and the other more fuel). Modern jet engines are extremely efficient at higher altitudes (FL300 and above) so you want to be there as fast as you can, irrespective of the sector.

@stali: The TO weight of long haul flight is way heavier than short hauls. Take again the B777 for example - TOW for VHHH-RCTP, full load is 495. Max TOW is 768. Unit is 1000 pounds.

@kevin Sorry to ask but how do you obtain TOW of 495k lbs.? Is it publicly available?


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Do negative flaps increase glide ratio?
If I understand correctly, gliders uses negative flaps while crossing between 2 thermals, but I don't understand why. For me:

positive flaps decrease sink rate but increase drag and thus worsen glide ratio
best L/G ratio is reached when flaps are in neutral position.
negative flaps decrease lift and thus require a greater angle of attack, maybe increasing drag.

The flaps on a glider wing are mainly for shifting the laminar bucket of the wing to the right lift coefficient. Yes, they help to increase maximum lift, too, but for gliding this is less relevant.
The goal in glider flying is to get in the shortest time from A to B. This can be achieved by flying faster than the best glide speed, because it gives the pilot more time to climb up in the next thermal. The stronger the thermal, the faster the optimum glide speed. You can read a great discussion about the finer points of this speed-to-fly theory here, courtesy of John Cochrane.
A fixed airfoil needs to have a wide laminar bucket, so drag can be minimised both when gliding and thermalling. This, however, will result in a less deep bucket which means that the drag coefficient within the bucket is higher than for a flap airfoil with optimum flap setting. Since adjusting the flap allows one to shift the laminar bucket to the desired lift coefficient, the laminar bucket of a flap airfoil can be made narrower and deeper.
The direct answer to your question is: no, negative flaps do not increase the maximum glide ratio but help to fly fast at low airfoil drag. If you look at the L/D at high speed / low lift coefficients, then yes, negative flaps improve the L/D at that part of the polar. They also help to limit the maximum lift a wing can produce and allow to fly faster in gusty weather.
Positive flaps reduce both drag at high lift coefficients and sink rate, and they are essential to reach the polar point of minimum sink in high performance gliders because that optimum is at very high lift coefficients due to the high aspect ratio of the wings. If you fly fast with positive flaps you will greatly increase drag and the sink rate, however.

Of course they do! Why would we want to use flaps if they wouldn't? The improvement in glide ratio happens, however, only for a certain airspeed range.
Look at the polar of the ASH 25:

I've added color to the polars to make it easier to visualize. The original polar was measured by DLR and Idaflieg and can be found in the flight manual.
For a given airspeed the glide ratio will increase with smaller sink speeds. If you're flying with 160km/h, your glide ratio will be better with a negative deflection (W(ölb)K(lappen)-Stellung 2) than without deflection (WK 3). To improve your gliding ratio at higher speeds you need an even higher negative deflection and should set WK 1.
The optimal flap position for each airspeed depends on your mass and is often given in the flight handbook

Source
The best glide ratio of the glider worsens with flap deflection (for both positive and negative deflections). This, as explained in Peter's answer, doesn't matter because the glider's cross country average speed depends also on the strength of the thermals. It is often better to fly with higher airspeeds than the best glide ratio airspeed.


text:
Why are magnetized materials dangerous goods?
I was reading about Dangerous Goods, and came across the fact that magnetized materials are classified as Dangerous Goods (item 12 in the linked page). We all see magnets every day, and they seem harmless.
Why are magnets dangerous to airplanes and require more careful handling?

@coderworks  Any reference on permanent magnets causing interference to pacemakers?

98% percent of magnets - the everyday ones - are completely harmless. I believe the concern is with super-strong rare-earth magnets, where they can be problematic. Aside from Simon's valid point on interference to compasses:

Getting stuck to things, such as other bags or the infrastructure itself, in the baggage system.
Attract themselves to other packages and content, possibly damaging these. Think of hard drives and electronics.
Being a nuisance and safety concern should somebody decide to open it, such as customs or security. You can get your hand crushed in the process if you open the box and the thing springs out.

Be aware they do not appear to be banned - just regulated in that they have to be properly labelled and packaged for safe transport. More curiously in that list are 'First Aid Kits' and 'Seatbelt Pretensioners'.

Depending on how exactly you interpret "First Aid Kit", it may contain flammable liquids (alcohol), pressurized canisters (e.g. spray-on bandages, inhalers), sharp objects (needles, scissors). That may be the reason, although I don't necessarily see how that makes it more dangerous than someone carrying his shaving kit (blade, compressed shaving cream, maybe rubbing alcohol) as baggage.

Some pre tensioners contains explosive squibs and/or gas cartridges.

Total/half-joking speculation: they don't want bad actors to have the tools to stay safer when they do their badness.

99% of magnets? =)

@mins while there are strong magnets within the hard drive, strong magnets + hard drive = Bad Things™.

@FreeMan: That's correct, deleting my comment... I did some research and degaussing devices used to wipe HD out are in the range of 1,5 tesla, not so huge compared to a strong magnet.

@FreeMan: Also, hard drive magnets are encased in mu-metal shielding to keep them from scrambling the data on the drive platters.

I use neodymium magnets in my work as a research scientist and I regularly carry them aboard aircraft in packaging.
We are not talking fridge magnets here, these are serious high-field permanent magnets. If two of them get properly stuck together they cannot be separated again easily. (The fields are far too strong to be pulled apart by hand and you can't use any metal tools for obvious reasons. One trick is to use a heavy, non-ferrous plastic-coated mallet to smash one off the other at the edge of a table, but even then it's usually half the day wasted.)
The problem is that the strength of the force between two such magnets is not intuitively understood by people unfamiliar with them: for example, your typical customs or security inspector. There might be almost no noticeable force between two magnets when they are some distance apart in the packing box but once they get within 5 to 10 centimetres they will suddenly be attracted together and anything in the way (flaps of skin, fingers) will simply be crushed flat under the force.
Or if nothing slows the magnets down as they come together, the force of the impact can chip the surface of the metallic protective layer, spraying shrapnel and metal fragments everywhere which poses a risk in particular to eyes. (The alloys used in constructing these magnets are extremely brittle.)
If the magnets get stuck to other large metal objects it can be a pain to separate them. They also induce weak magnetism in metals you normally expect to be non-magnetic.
Finally if they get wiped across anything like a credit card or ID badge that's usually the end of it (not always actually, it's a bit random).
The magnets are not extremely dangerous, getting small bones broken or metal fragments in the eye is the worse case scenario on my risk assessment, but they are not toys and there is a risk to those untrained in handling them. So anything like this needs to be identified so that they can be treated with care if someone does want to open my luggage to fish around inside it.

Could two of those magnets be safely sandwiched on opposite sides of, say, a thick slab of very dense wood?

@Sean Depends what you mean by "safely". They are safe until they are disturbed and handled by someone who isn't trained to use them. They were usually wrapped in bubblewrap individually and packaged something like the way you suggest. (I no longer work in research, so my information here might be out of date.)

Because the aircraft has a standby compass in the cockpit which is a normal magnetic compass.  This compass must be fitted and must be working in case of extreme instrument failure.  There are three things which standby instruments must provide; height, speed and heading (the compass).
Magnetic cargo can interfere with this compass and therefore, is treated as dangerous cargo.  Whilst the risk is low, the aviation industry leans on the side of safety, and we can all see the benefits of that.
This report will tell you what you want to know.

"Magnetic field decreases exponentially with an increase in distance
as measured from the compass." (from the report linked above). Really? I thought for a typical permanent magnet it decreased with an inverse-cube law - faster than inverse-square, but not "exponential". Inverse cube would give a reduction of about x10 between 2.1m and 4.6m distance, which is roughly consistent with Appendix A of the report.

@alephzero Never was a whiz at maths, but surely an inverse cube is exponential?  Certainly when plotted, it looks exponential to me.  Regardless, rather than say "no magnets, unless in a metal container at least 10m from the cockpit" take the safe route and classify magnetic material as dangerous.

FYI "The graph illustrates how exponential growth (green) surpasses both linear (red) and cubic (blue) growth."

Well, I would hope ICAO get the math and physics correct in their official documents. "Exponential" means the rate of decrease at any point is proportional to the strength of the field itself. That means it would decay faster than any "inverse power" relation. The field from any geometrical arrangement of permanent magnets can only decay as some inverse power. For special cases like a "quadrupole magnet" it could be faster than an inverse cube law, but for this risk analysis the slowest possible decay of the field strength is what matters.

@alephzero: From a practical perspective, exponential should be classified as neither faster nor slower than quadratic or cubic.  If some quantity decreases quadratically with distance, doubling the distance will cut it in half, regardless of whether the distance is large or small.  If some quantity drops by half for each mile of distance (exponential decrease), doubling a ten-mile distance would reduce it a thousandfold, but doubling a ten-foot distance would hardly decrease it at all.

@supercat " If some quantity decreases quadratically with distance, doubling the distance will cut it in half" - half??? ;)

@alephzero: Oops.. Factor of four.  My point was that the ratio by which doubling the distance reduces a quadratically-decreasing quantity is independent of the initial distance, while the ratio by which doubling the distance reduces an exponentially-decreasing quantity is totally dependent upon the initial distance.

@alephzero I thought it was inverse fourth, actually.

@Random832 It is possible to construct an array of magnets or shape magnets to have quite a complex field geometry locally. But once you are sufficiently far away the inverse cubic term dominates the decay rate.

Airplanes have many sensors which are sensitive to electrical noise. And moving magnets can cause an electrical flow, hence electrical noise. As you can not estimate who is carrying how strong magnets it is safer to forbid them in life-critical environments like airplanes.
And about compass, it is only needed for an initial heading estimate, and it replaces GPS heading after moving a few meters in aircraft. So the compass is not a big deal as long as GPS is available.

Do you have a source for your last paragraph? I'm not sure why they would go to the trouble of calibrating the compass if it's "not a big deal as long as GPS is available."

Sorry for misunderstanding. I was trying to say commerical airplanes uses latest technologies and not solely relying on compass measurements in the air. But the risk is u can not know gps will be availabile all the time. Safety and redundency issues. Compass is used for IMU (Inertial Measurement Unit) initialization/alignment on the ground pre-takeoff. And it depends on the INS technology and how it uses compass measurements (How it contributes to the kalman filter).


text:
Where can I get a reliable and a updated Airport API or Airport Database, and I also want to know the charges for the API?
Kindly guide me where can I get a reliable and updated airport Master API. The airport API should have all the information about all the airports in the world and I also want to know how much do they charge for the API.

what does how would be the plans and pricing for the Airport master mean?

I just want to know where can i get the airport API and how much does it costs.

I've voted to reopen this question.  It's clear what the person is asking for, the fact that it's out of his price range doesn't mean the question can't help future visitors.

I'm voting to close this question as off-topic because resource location is specifically off-topic defined in the help center.

@gerry do you think it would be better suited in opendata.SE

@ManuH I believe it would.

Location Identifiers & Airline Designators data files available with daily updates
The perfect solution to feeding your coding database. These files (more information in right hand column) are available as a 12 month subscription or a one-time web download. For the subscription, you will be granted access to the password protected extranet site upon purchase where you can retrieve the data files update daily.

http://www.iata.org/publications/Pages/coding.aspx

Thanks for the information RedGrittyBrick, but the price for the data is to high are there any other sites where i could get the data at a lower cost.Since we are a startup company we cant afford such high price.

@NivinAbraham If you can't afford IATA, check out ourairports.org. Also see http://aviation.stackexchange.com/questions/2644/can-two-airports-have-a-different-icao-code-but-share-the-same-iata-code

Thanks a lot Danny Bekett, the URL you provided was very helpful.The data are  very accurate and the also it is up to date.

Navigraph and Aerosoft provide relatively cheap (~€25/yr) subscription to regular AIRAC updates for the simulation community. The databases include all the airport information that is required for FMS (navigation).
In both cases, the databases are plain-text files which are rather trivial to parse without any special API. For example:

A,YMML,MELBOURNE INTL,-37.673333,144.843333,434,10000,11000,11900,0
R,09,83,7500,148,0,0.000,0,-37.660756,144.822247,395,3.00,50,1,0
R,16,161,11998,197,1,109.700,160,-37.653192,144.834911,432,3.00,50,1,0
R,27,263,7500,148,1,109.300,263,-37.662261,144.848089,407,3.00,50,1,0
R,34,341,11998,197,0,0.000,0,-37.685789,144.840994,330,3.00,50,1,0

IMS,YMML ILS/CAT III,109.700,0,1,195,-37.656667,144.833611,434,YM,0,0,11.5,0.0
IMW,YMML ILS/CAT I,109.300,0,0,195,-37.660000,144.810000,434,YM,0,0,0.0,0.0

One can get a free sample of the whole database (at some point in the past) with a demo version of some software packages that use such database; for example, this one.


text:
Why are geodesic airframes no longer produced?
Why are geodesic airframes no longer produced?  Do they cost more than conventional structures?  Are they not as strong as conventional structures?  I assume it is insufficient cost benefit ratio, but how does that work out?
 Wellingtons under construction. Photo: Royal Air Force

Probably because the skin will take a lot of the load. especially in a pressurized craft

A search for "isogrid & aircraft" shows interesting (including current) applications. Isogrid structures are considered for composite airframes too, but at the moment they are not cost-efficient compared to state-of-the art manufacturing concepts.

Geodesic (sometimes also spelled geodetic) designs were used in aviation as early as 1909, in the Schütte-Lanz SL1 airship. It had a wooden structure with fabric covering, and Professor Johann Schütte, the scientific head of this design, used the most efficient method conceivable. However, planning the shape of all structural members was an enormous amount of work, and stretching an existing ship by inserting a new section was almost impossible. For those reasons, the late Schütte-Lanz airships used aluminum frames with conventional design, just like Zeppelin did from the beginning. Diagonal wire bracing was used to take up shear loads.

Geodesic frames are good in transferring both bending and shear loads. This is helpful when the aerodynamic covering is not stiff enough to contribute any shear stiffness, as in the case of fabric covering. However, when the skin is made from the same material as the frame, it can carry shear loads and a geodesic frame would not improve matters. Now it is better to carry longitudinal loads in longerons and shear in the skin.
But geodesic constructions are not dead. CAD makes the design effort manageable, and in some cases they give a weight advantage. Look at the casing of the EJ200 engine: It uses a geodesic reinforcement to make the casing light and strong and to prevent harmful vibrations. In the end, a jet engine casing is a big pressure vessel and needs to be strong, stiff and light.

photo by Julian Herzog, CC-BY

The switch from internal frames to structural skin airframes -- called monocoque construction -- is fairly well covered in Wikipedia:

Early aircraft were constructed using internal frames, typically of
  wood or steel tubing, which were then covered (or skinned) with
  fabric3 such as irish linen or cotton. The skin added nothing to
  the structural strength of the airframe and was essentially dead
  weight beyond providing a smooth sealed surface. By thinking of the
  airframe as a whole, and not just the sum of its parts, it made sense
  to adopt a monocoque structure and it did not take long for various
  companies to adopt practices from the boat industry such as laminating
  thin strips of wood

That being said, we are still producing tube and fabric airplanes, particularly taildraggers based on classic designs, such as this American Champion Citabria and this Super Cub:

dat little Cubbie's got such a cute face!

It started when the usual way of building was a steel-tube frame, with wooden formers attached to carry the skin.
Early geodetic planes were still fabric covered, but the entire airframe structure was now in the skin, leaving the interior open for payload like people or ordnance. It was much lighter and stronger, though patented and many builders simply don't want to license-build as much of their plane from somebody elses' ideas.
Once metal skins were being used, it wasn't the only monocoque option, though it was still superior to the usual orthogonal structures. Still more expensive though.
"Geodesic" should refer to the field of mathematics itself. Geodetic applications of it work for space-frame structures or mapping a globe, etc.

Geodesic airframes are useful for purpose-built crop dusters, built from 4130 stainless steel tubes. The chemicals on board could easily react with aluminium, which is still used for skin panels but not as a structural member: the skin plates are only for covering, not for taking any of the prime loads. The construction method is one of the reasons why they look unlike any other plane, with the angular cockpit shape and all.
Other than that, monocoque construction has completely taken over from geodesic frame construction - when the skin plates are the main load bearers the airframe is lighter.

Is that the only reason they look different? I would think the requirement for good visibility in low-level flying is pretty important.

I was under the impression that the odd shape of dusters had more to do with the need to keep the CG forward by inserting the chemical tank between the engine and the cockpit.

You say "Crop dusters have..." as a fairly blanket statement covering all crop dusters. Would that be better stated as "Some crop dusters have..."?

@jamesqf - you mention visibility. To me (IANAP) the visibility out of that looks terrible with that very long hood and the pilot sitting directly over the wing, though I'm willing to be convinced that I'm very wrong.

As other answers already note, the major attraction of geodetic framing was to provide a lighter and stronger structure than internal framing, with the not inconsiderable bonus of leaving much more internal space for fuel and payload.  The geodetic frame could directly carry a fabric skin in a streamlined shape, instead of requiring the addition of (heavy) falsework merely to carry the skin.
Some modern aircraft are still built in a frame-and-fabric manner, such as the Aviat Husky, though most use a more generic "space frame" construction rather than geodetic framing.  These are almost universally General Aviation aircraft designed principally for VFR flight.
A general disadvantage of fabric-skinned aircraft is that they do not easily support cabin pressurisation; it would be necessary to install a second, internal pressure hull to retain a sufficient air seal.  While this would not be especially difficult to achieve, it does reduce the main advantage of the geodetic frame, in the context of aircraft designed to regularly fly above about FL120 (where most people start to experience debilitating hypoxia) without carrying oxygen for every occupant for the entire cruise phase of the flight.
Instead, metal-skinned and composite aircraft are typically designed so that the skin forms the pressure cabin as well as the aerodynamic surface, and also carries some of the structural loads so that less internal framing is required.

One area where geodesic airframes have seen a recent resurgence, is 3-d-printed model airplanes.
I've seen an example where the fuselage and wings and tail surfaces were all of open geodesic construction, similar to the well-known Vickers "Wellington".  The model was covered by a heat-shrunk plastic film such as Monokote.
Here's a website featuring many examples of such models-- https://www.3dprintedrcplanes.com/
This link describes the process of designing and 3-d-printing double-walled structures for model airplanes, with an internal geodesic structure between the inner and outer walls-- https://www.rc3dprint.com/post/3d-printing-geodesic-internal-structure

They ARE still produced. See for example the EU research project from 2013 here: http://www.transport-research.info/project/advanced-lattice-structures-composite-airframes

Welcome to Aviation.SE! Please consider adding more details to your answer, to make it more qualitative.

There is an interesting PDF if one searches a bit on the linked page. However, it seems the subject is research, not yet production.

Link no longer works.


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How does the IAS stall speed vary with increasing altitude?
First of all, we all know what IAS is and if not for some people you can follow this link to read the basics about stall and the differents speeds. So we also know that our TAS/ Mach number will increase with increasing altitude.
Let's assume that our IAS stall speed is 100kts at sea level, what will be our IAS stall speed at 30 000 feet with exactly the same aircraft ? (Assume ISA atmosphere)
Everyone would be tempted to answer that it doesn't change, right? but what about the low speed buffet in this picture ?
We agree in saying that the low speed buffet here is not the same as the one at sea level but look, our IAS/CAS stall speed has a higher value now. Maybe that it's a representative value of IAS/CAS calculated by the Air Data Computer from an EAS but still, our IAS stall speed has risen.
We can also keep it simple and think about it for a non pressurized single engine aircraft. I don't need an exact value but I would love to have a logic and smart answer.
It seems to be very basic but when we correctly think about it we can find ourself a little bit confused.
For me, the answer is : The IAS stall speed will increases

"IAS stall speed is 100kts at sea level".  What do you mean?  An aircraft can stall at any speed.

I know that, I know that an aircraft stall at the same angle of attack regardless of its speed. I just took an example

@Simon I think it's safe to assume he is referring to the straight and level flight stalling speed. This type of question is asked in various theory publications and I'm not sure why it has a close vote

That's visible on the coffin corner diagram. Also on Wikipedia (U2 diagram).

Airplanes do not stall at the same indicated speed or even at the same angle of attack - it all depends on circumstances.
The angle of attack dependency is discussed here. An increased pitch rate can push the stall angle of attack 50% higher than what the stall angle of attack is in stationary conditions.
The next big factor is the Mach number. When increasing the angle of attack, the flow around the airfoil's nose will develop a suction peak. This suction is equivalent to higher local speed, and if the critical speed (when local flow speed equals the local speed of sound) is exceeded, the flow past the suction peak will no longer behave similarly to the flow at the same angle of attack but a lower flight Mach number. Let's just say that the local Mach number in the suction peak has a strong influence on the stall angle of attack, and flying at a higher Mach number lowers the stall angle of attack, sometimes dramatically.
Increasing altitude will raise the flight Mach number in two ways:

The reducing density means you need to speed up to fly at the same dynamic pressure, and
the atmospheric lapse rate decreases the speed of sound in air.

Both effects conspire to reduce the stall angle of attack at 30.000 ft to a value quite a bit below that at sea level. Details depend on the airfoil and specifically on its nose radius and wing loading.
Only very light aircraft will not be affected by the change in Mach number, but even here the stall angle of attack at altitude is lower than at sea level due to the reduction of the flow's Reynolds number with increasing altitude.
In short, the indicated stall speed goes up with increasing altitude for a variety of reasons and does so nonlinearly. The magnitude of the change depends on a multitude of details.

Thank you for this precise answer! I waited for this type of answer from you specially! Can you just tell me what would you answer to my question ?

@Thomas: To answer the question I need much more information. Airfoil and wing loading would be the most important details, but maybe not sufficient. I tried to explain that the change in stall angle of attack depends on many things, and your question only gives a sea level stall speed.


text:
What are the airflow patterns of a helicopter when operating in ground effect?
When operating in the air cushion region of approximately 8 inches and below to the ground what pattern does the downward flow of air flow?

What happens at a height of 8 inches and below in the hover in ground effect is a function of rotor diameter: everything is relative. Ground effect becomes noticeable at about one rotor diameter from the ground surface, so if your rotor is 8 inches in diameter there is not yet a significant change of the downward air flow.

If we limit ourselves to discussions of flow as a function of rotor diameter, we can see the following happen as illustrated in the above figure, from J. Gordon Leishman, Principles of Helicopter Aerodynamics:

Out of Ground Effect, the ground has little or no effect on the rotor outflow. The spiralling lines in the figure are the blade tip vortices.
In Ground Effect, the rotor downwash appears compressed and spread out. One way of understanding how ground effect affects lift, is by considering that the static pressure underneath is increased and contributes to the lift.

The spiralling blade tip vortices are compressed and spread out as well, and this has a large effect on wake fluctuations, particularly noticeable for a ground observer. What exactly happens regarding spiralling gusts is not yet fully understood - from Prouty, Helicopter Performance, Stability and Control, page 68:

The wake in ground effect is not really a steady flow..., but a flow with large scale fluctuations that can be felt as gusts by an observer standing near the helicopter. Reference 1.21 speculates that these fluctuations are associated with the vortex that is made up of the individual vortices from the blade roots.This vortex apparently writhes like a pinned snake along the ground, causing the entire wake to shift and wobble.
Pilots occasionally report that when hovering in ground effect, they experience random yaw disturbances. This is probably due  to the effect of the root vortex as it writhes near the tail rotor.

Thanks for the input Koyovis! I had heard that many people speculate that the pattern of the downwash takes on a horseshoe pattern have you heard anything on that?

No I have never seen a reference to horseshoe shaped downwash.

I was going to fix this sentence, but I'm not sure what it's supposed to be: and this has a large effect on was fluctuations Is was supposed to be wake?

Thanks @TomMcW, sometimes autocorrect makes weird autocorrections. The original splelling mistake was of course mine.

Well, once the rotor is moving in one direction of course will a horseshoe vortex result, like with a wing. Only with a more complex flow pattern.

Koyovis you commented on a similar question that I posted, “How to keep a helicopter steady when close to the ground,” and mentioned how people would feel large fluctuations of air when standing next to the helicopter. That example coupled with how the root vortices cause random yaw movements for pilots makes perfect sense. My question is whether or not we can anticipate these movements to counter them?

@Wes According to the description from Prouty the disturbances are of a random nature, which would imply that timing and magnitude cannot be anticipated. I could not find any frequency distribution plots of the rotor down wake affecting the vertical tail - it could be that some of the vortex effects are of a periodic nature because of the blade passing frequency, but I could not confirm this.

@koyovis I greatly appreciate taking the time to look for the answer to my question! Unfortunately like you stated, these disturbances cannot be anticipated but I would like to think as you do that there would be some what of a consistent pattern due to  the vortex effects of the blade passing frequency.


text:
Why are there fewer stages in the turbine than in the compressor of a turbine engine?
I'm a student interested in aviation. I'd like to understand why the turbine section has fewer stages than the compressor section in a jet engine.
What would be the effects of adding stages to the turbine, or to the compressor? How is the optimal number of stages determined?

What do you mean "less than"? Less than what? Compression ratio's?

Are you asking why there are fewer turbine stages than compressor stages? E.g. 2 turbine stages but 10 compressor stages in many high pressure systems.

Because the hot side has a larger volume and energy and doesn't has to be very effective to harvest the energy it needs. Compared with the compression ratio of the compressor, the turbine doesn't decompress as much .

Very simply, a turbine has fewer stages than the compressor, because the flow in the turbine is going in the direction of reducing pressure - the air wants to go that way naturally. But a compressor has to make the air flow in the direction of increasing pressure. If this is done too aggressively (attempting too much pressure rise per stage), the air flow on the blade seperates, and stalls. In severe situations, the flow can completely reverse direction and exit the intake.

It has to be easier for the gas to go through the turbine than back the compressor, otherwise it wouldn't flow in the desired direction.

It's well-known that the designers always desire to less weight and less complexity, so if there isn't the need to add the stages to the turbine they will not add auxilary stages. Such, the problem is why the compressor need much more stages to achieve some compression ratio  than the number of stages of the turbine,simply to say, with somewhat less (numerically) expanding ratio. If exactly, the expanding process is continued in the exhaust nozzle. But this isn't the main cause of the less number of stages of a turbine.
The explanation is following.
Every stage has a definite max loading coefficient that determinaned by aerodynamical possibility (certainly, w/o significant losses) to change flow direction on a stage (of a turbine or a compressor). The aerodynamical possibility of a hot gas flow with very high temperature (i.e. with very high sound speed and very less density, conseequently,  the flow is subsonic and with high Re number ) is much more than the possibility of the cold gas flow with  higher density  on a compressor, because the cold flow is with lower sound speed, and usually the flow is supersonic or high subsonic, and futhermore with much lower Re number (more influence of viscousity). It causes great difficulties for significant change of flow direction w/o essential losses.
And it's necessary to add a very important issue that the gas flow in the compressor moves towards the zones with higher pressure (pressure gradient is positive) that may lead to flow separation (contrary flow in boundary layer), while in the turbine the negative pressure gradient exists that is prevent factor for flow separation.
That is why the turbine blades are made thicker with great curvature and more "streamline" than the compressor blades.
Therefore, a turbine
stage has much more loading coefficient than a compressor stage and conseequently a turbine may be made with less number of stages.

Well done - a good effort for an answer. I think the gist of what you are saying is correct although the answer is difficult to read at times. Although for such a complex design space it is difficult to explain in just a few paragraphs. As you say the aerodynamic limitations means that compressors are limited to about 1.4 pressure ratio per stage so you need multiple stages to get the required compression for the performance. Turbines extracting energy can be impulse or reaction design and only need one or two stages to extract the required energy.

From windmill physics we see that wind power is proportional to the cube of windspeed as:

Wind power = Density × Area × v$^3$/2

The exhaust gasses of a jet engine are of lower density from heating, slightly higher mass, but are going much faster than the inlet gasses
This explains why a smaller turbine can run a larger (more blade area compresser) with plenty of energy left over for propulsion.

How to use excess energy?

The most efficient method of propulsion depends on the desired airspeed.
At low airspeeds propellers (literally rotating wings) are most efficient at generating thrust.  Energy is extracted to turn them by adding more turbine blade area.  This reduces exhaust gas velocity but extracts more mechanical energy.
At higher airspeeds the fan replaces the propeller.  At transsonic and supersonic speeds, the "extra windmill" can be removed from the turbine and the engine operates as a "turbojet", essentially an airbreathing rocket with the turbine only running the compresser.
Faster still, one can remove the compresser and turbine and operate as a ramjet.


text:
How does the raked wingtip of the Boeing 787 work?
The wing tips are very different on these airplanes. The wingtip on a Boeing 787 is a sharp triangle, while the wingtip on a Boeing 737-300 is flat.

British Airways Boeing 787-8 G-ZBJC wingtip | by StephenG88 (source)
Why is the wingtip on the Boeing 787 different from that of the Boeing 737?

I promise you guys a picture of what I mean but my laptop has not been working correctly

possible duplicate of What do winglets do to increase aircraft performance?

@Simon: No, Ethan refers to what is called a raked wingtip. Works better than a winglet, too.

@Simon I have to agree with Peter Kämpf, raked wings work differently than winglets (even though they do similar things), so this is not a duplicate of that particular question.  And none of the answers really dive into how a raked wings works so...  Yeah, not a duplicate.

@PeterKämpf At the time of my comment, the picture was not in the post and there was only mention of a "triangle" which I took to mean a sharklet.  Got it now so removed my vote to close.

Guys I was asking about the shape of the wing tip not about the 787 wing and the drag reduction of the wing tip vortex thanks some of those pictures were helpful. I have figured out the answer on my own, but thanks for putting effort and research into my questions.

The purpose of the raked wingtip is same as that of the winglet- to reduce drag by reducing the effect of the wingtip vortices.

Source: www.boeing.com
Most aircrafts use winglets to improve performance. Basically, they reduce the wingtip vortex by [deflecting the air](and the associated drag by deflecting the air that wants to escape over the tip back down) that wants to escape over the tip back down. This reduces the induced drag.
In case of raked wingtips, the tips of the wings have more sweep compared to the rest of the wing.

"B787-800v1.0" by Julien.scavini - Own work. Licensed under CC BY-SA 3.0 via Commons.
The raked wingtips also reduce the wingtip  vortex- by redirecting the wingtip vortices further outboard and aft of the wing.
The raked wingtips are more efficient compared to the winglets in cruise and are used in ultra long range aircrafts.

Source: www.boeing.com
One problem with the raked wingtips is that they increase the span. This increases the bending moment. Another reason we are not seeing a number of aircrafts with raked wingtips is due to the fact that they can't be retrofitted (Though there was a proposal to retrofit KC135 with raked wingtips).
Note: The photo is not very clear. I'd simply answered about the raked wingtip.

I am afraid I will never understand how a vortex behind the wing can cause drag.

This is due to the three dimensional nature of the flow; if the flow was in two dimensions (i.e. around an airfoil), there would be no wingtip vortices and no induced drag.

@aeroalias, no, Peter is right. The vortices can't cause drag. They have the same cause as induced drag, but they are not the cause themselves.

The wingtip vortices cause the downwash, which reduces the effective angle of attack of the wing. This effectively tilts the lift vector backwards, causing a downstream facing force (drag) on the wing. This drag is called the induced drag as it has been induced by the action of tip vortices.

Yes, and wet streets cause rain. Everybody knows this -it's obvious!

@PeterKämpf: Drag is a continuous loss of energy. In particular, it's mechanical energy lost by the plane to the air. Some drag is unavoidable, associated with the forces that bed air downwards to generate lift. Forces that accelerate air in other directions still cause drag, but no lift. The vortex is proof of air moving in unintended directions, and thus of avoidable drag, but the vortex indeed is not the force itself. It may appear counterintuitive, but physicists do this all the time. Ask CERN: all they do is looking at the debris of a collision to say something about a particle.

@MSalters OK, I'll bite. This is about cause and effect, and aeroalias has it backwards (as do most people here). Or do you mean that baryons in CERN cause collisions? Certainly not, but this is what aeroalias says about the wake vortex. The vortex is proof of lift and as such is unavoidable. It is just what happens behind the airplane to even out the disturbance it has caused, but it does not cause drag. If you read some more of my answers, the cause-effect relationship should become clear.

@PeterKämpf: What I was saying is that a lot of physicists aren't considering cause and effect in such black and white terms. I agree that both the drag and the vortex are associated with the acceleration field of air flowing over and under the wing. However, the vortex is part of that acceleration field, while the drag is caused by the field.

@MSalters: I prefer a clear understanding of a working mechanism to a muddled view where cause and effect are interchangeable at will. We have a clear sequence of events because airflow and time go in one direction. Even at CERN time does not run backwards. Why then accept such nonsense like "The wingtip vortices cause the downwash"?

@PeterKämpf: Looking at airflow and time go in one direction is a rather complicated approach. A physically equally valid view is to study it as a steady-state model. This freezes time out of the equation, making the problem easier to solve. This is in fact a common technique in physics.

The raked wingtip of the Boeing 787 is now "le denier cri" in wingtip design, but has been successfully employed by water fowl for many millennia. The use of composite structures has made the three-dimensional shaping of wingtips much easier, and the first airplanes to use increased sweep at the wingtip were indeed composite gliders such as the Schempp-Hirth Discus.

Seagull (left) Schempp-Hirth Discus 2 (right, source)
In the times of metal wings, Dornier introduced their TNT (Tragflügel Neuer Technologie, in English: Wing of new technology) on aircraft like the Do 228, Do 328 and the Dornier ATT flying boat. It used a triangular wingtip with no change in trailing edge sweep, which is easier to manufacture in aluminium.

An Air Alps Dornier 328-110 landing at Fiumicino Airport, Rome, Italy, viewed from Ostia Antica. Picture by Makaristos.
You can think of the raked wingtip as of a winglet which has been folded down. Increasing the wingspan will involve a bigger mass of air into the creation of lift, and the induced drag is reduced. This explains already most of the effect, but the increased sweep adds some more benefits.

Higher sweep lowers the lift curve slope. This helps to reduce the lift contribution (and, consequently, the bending moment!) of the wingtip at high angles of attack. Therefore, raking the wingtip will result in a lower maximum bending moment.
The sweep also shifts the center of pressure of the wingtip backwards, adding a negative torsion moment. This moment lowers the angle of attack over the whole outer wing in gusts and at high load factors, again helping to lower the maximum bending moment at the root.
Especially in gliders, it allows to combine a near-elliptical circulation distribution over a wide range of angles of attack with good stall characteristics. The lower lift curve slope of the wingtip lets it stall at a higher angle of attack and reduces its lift contribution relative to its area at increasing angle of attack.

This aeroelastic tailoring, together with the low fuel consumption of modern high-bypass-ratio engines, allows the Boeing 787 to use a wing of an aspect ratio of 10 and the A350 one of up to 11.8. Previously, - apart from gliders - only special designs for extreme flying time and range like the B-24 Liberator or the Breguet Atlantique sported such high aspect ratios. Older airliner designs had to use lower aspect ratios (between 6 and 8.5) to have enough wing volume for fuel.


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Would an engine have more thrust if it was mounted on top of the wing?
NASA actually did studies that say that it does increase the thrust. Air moves faster on top of the wing than the bottom, so could this have an effect on the engine's performance if there is more intake of air into the engine if the engine was mounted on top of the wing instead of the bottom?

How about posting a reference to these NASA studies?

@Ethan, questions don't need images. If you refer to a study, you should include a reference to that though.

I doubt anyone here is going to disprove a NASA study but you really need to get a handle on what exactly you're asking. If an engine is getting the same airflow and fuel then it's going to produce the same thrust no matter where it is - but there is far more than that at play when it comes to the effect it has on flight.

@Dan I am specifically asking if the thrust would be higher if the engine was on top of the wing because the air flows faster on top of the wing than on the bottom, so the more air received by the engine. Sorry for the confusion

@mezzanaccio I apologize I could not find a link for this I just remember reading a article about the studies of the engine on top the wing and how it could increase the thrust. I will do a little more research for you.

First, we need to agree what thrust is. The simplest definition would be to measure the pressure in the intake and the nozzle, as described in the answers to this question.
But this gets you only so far. It would be more complete to use the difference in horizontal forces acting on the plane between idle and full thrust for the definition, because that is really what interests the pilot. Imagine that your engine will cause massive separation around the intake at idle (spill drag), or that the exhaust jet will help the flow over the deflected flaps to stay attached, avoiding the drag of separated flow.
If we can agree to include the change in drag into our thrust measurement, the next question is what flight segment is used? In cruise, the flow around wings and fuselage is nearly completely attached, and the engine can only spoil the orderly flow. Next, since the speed of flow is higher on the upper side of the wing and the intake needs to decelerate the flow, a top-mounted engine would need to decelerate its intake flow more than one below the wing. This leads to higher intake losses and less net thrust. Conversely, if the engine is below the wing, the wing will already decelerate the flow somewhat before it is ingested by the intake, which makes the work of the intake easier and reduces losses.
In the landing configuration, however, a deflected flap blown by the engine exhaust will create a much better L/D than one without. In this case, it is better placed on top of the wing. The Boeing YC-14 and the Antonov 72 are examples of such configurations.
Since the cruise segment is much longer, most designs prefer to place the engines below the wing, but quite a bit ahead of the aerodynamically ideal position to shift the center of gravity of the wing ahead of its axis of elasticity, thereby helping to dampen flutter.


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How can you calculate the gliding distance for a remote controlled jet?
Is there a way to calculate the gliding distance for a remote controlled (RD) jet aircraft, in case of engine failure? We have some telemetry data on the model, giving us speed and altitude so as to plug them on an equation or other model in estimating the glide distance, but we would not want to actually risk destroying the jet by testing a real life scenario.
So, for example, our maximum allowed altitude is 500ft Above Ground Level and its speed is about 150 knots; we are interested in calculating the gliding distance if engine power is lost or cut off (the engine is automatically cut off if the jet flies outside radio control range).

You need to be a little more specific - what is the aspect ratio, wing shape, fuselage size etc. This would allow to estimate the drag, which is a necessary part of calculating lift/drag. This ratio will give you the glide ratio.

I don't see the risk in testing. Set it up ok a nice day, cut the engines, measure the glide, and restart the engines at, say, 100ft AGL.

Hi Peter, thanks for your answer. Assuming you have the quantities described above, it the computation straight forward or does it involve numerical analysis? I remember back in my grad school days we used Galerkin procedure to solve a number of differential equations, but it completely escapes me now.....is that what you are referring to?

I think the only way is to test it.
Choose a day with very calm wind.
Set the plane up for fixed speed anywhere between obviously-lower-than-best-glide-speed and obviously-higher-than-best-glide-speed, and make multiple tests with incrementing the speed in small steps (like 5kt). You can probably think out the details of the experiment yourself, but this should work well.

It's very difficult to compute these things. Companies like Airbus or Boeing have (probably proprietary) software for this, large clusters, etc., perform test in an aerodynamic tunnel and do many test flights. Since you likely don't have all the equipment, the only thing you can rely on is the test flight.

thanks for your response; I believe that the best way to measure it is through tests (using max and min speed reference boundaries as you indicate).


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Where do Russian and Chinese airliner manufacturers test their airliners?
Airbus test their plane on Toulouse Airport, and Boeing test on Boeing field.
How about Russian airliners Ilyushin & Tupolev and Chinese Airliners Comac ?

Note, that this only applies to the basic tests. In the later phases of testing the aircraft needs to fly to windy places, cold places, hot places and so on, which generally takes it all over the world.

Almost all Russian (Soviet) aircrafts were tested in the Zhukovsky Airfield (Ramenskoye Airport) near Moscow.
The Comac is building a Flight Test center at Dongying, Shandong province near the Dongying Shengli Airport. The flight test of aircrafts are expected to be carried out from here.
As Peter noted, Antonov test their aircrafts at Kiev Airport, including An 225, the largest aircraft in world.

Antonov is doing their testing at an airport northwest of Kiev, and did this even in Soviet times.

You are right. In my defense, Antonov is Ukrainian, not Russian.

… and they build few civilian airliners. I didn't mean to criticize your answer, just add to it.

I didn't take it that way. I was using a loophole :). Anyway, edited.


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Why is the combustor volume of a jet engine decided by altitude?
I have read that the combustor volume of a jet engine is decided by the maximum altitude up to which that aircraft will fly. I'm getting confused by that, because the maximum mass flow rate will occur at sea level so the combustor volume should be decided based on sea level. Can anyone explain how the volume is decided?

Are you sure it's about measuring the volume? The volume of a combustor is constant. But the higher you fly, the larger the engine must be. So, the max. altitude determines what volume is needed.

@sweber Did you misread the question? The question seems to be saying exactly what you suggest: that the design of the combustor depends on the desired maximum altitude.

I edited the question for clarity but I hope I didn't confuse things. I understood it to mean: "combustor volume is determined by the intended service ceiling of the aircraft, but maximum combustor volume is required at sea level, so if the combustor is big enough for sea level then it's big enough for any altitude and therefore the service ceiling shouldn't matter anyway". But since I know nothing about jet engines I didn't want to edit it so radically. Maybe the OP can comment.

Short Answer
It is less the volume than the length which limits the maximum altitude. The cross section of the combustor is determined by the volume flow coming from the compressor and the desired flow speed in the combustor, so this dimension is given. To enable operation at lower pressure, the length of the combustor must be sufficient, so a longer combustor will have a higher volume. That is why you have read that volume determines maximum operating altitude.
Background
The factors which make ignition easier are

pressure
temperature, and
spark energy of the ignition system

Ignition is made harder by

mean fuel droplet size (bigger droplets take longer to evaporate)
flow speed in the combustor, and
higher turbulence intensity.

For combustion to take place, first the fuel droplets injected into the air stream need to evaporate. This is made easier by higher temperature and pressure, and the lower both are, the more residence time of the fuel-air mixture in the combustor is required for good combustion. Longer combustors have higher pressure losses and weigh more, so engine designers try to limit their length.
With increasing altitude, the pressure in the combustor drops because the compressor will simply lift the pressure at it's inlet by a given multiple. Therefore, combustor pressure drops in sync with ambient pressure, and the required length for sufficient combustion goes up.


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How to turn a fighter jet?
I am not a pilot - just a student & play Ace Combat  sometimes
So I turned off the flight turn assistant in the Ace Combat flight simulator/game & now I have to manually control bank & pitch to turn. I also like to use a little yaw for stabilizing myself.
If I am not wrong - both pitching & rolling are controlled by the control yoke in jets. So, when going for a right turn from level flight, what does the pilot do:

Push the yoke a little right (attain 90+ degree rightward roll) and then pull the yoke towards him (to pitch up - i.e. actually execute the right turning arc). OR
Pull the yoke in a diagonal way i.e. towards SouthEast (assuming pilot was initially heading north) or 4'o clock - and the flight's internal computers break down yoke movement movement into respective pitch and roll magnitudes.

Just curious.

What is the purpose of the turn?  That is going to determine how the turn is executed.  Are you simply making a heading change, or are you executing a level break turn, etc?

@RhinoDriver no idea about level break turn just wanted to roam around, I rather let the enemy find me :)

if that's the case then you should look to the answer given.  This is less specific to fighters and more specific to basic airwork.

@RhinoDriver ok got it. Also, can you point me to pages describing fighter jet moves like level break turn, etc.

I think you are making things too complicated on yourself. To the best of my knowledge, fighter jet controls will in principle behave the same way as almost any other aircraft- they will in general actuate surfaces the same regardless of orientation. A left/right movement controls roll, a backwards/forwards movement controls pitch.

You would push the stick sufficiently to the right to bank until you reach the desired turn rate. You would have to maintain a little right stick to maintain the bank angle.
As you do the above, pull back a little to ensure that the vertical speed remains zero. You also need to add a little rudder to create a coordinated turn.

See the following images taken from here and here:

Thanks sir, just one more thing - in cars, after a right turn is executed the steering wheel is brought back to its neutral/zero position by necessary counter-clockwise turns of steering & sometimes the steering comes back to neutral by itself -- so what happens in planes?

This behavior depends on the aircraft and flight conditions. As a rough rule of thumb, since the wing turning inwards is moving somewhat faster, it will generate a little bit more lift and will normally cause the aircraft to gradually bank even. This can be further encouraged by for instance applying a dihedral to the wing. The exception to this rule are Airbus that maintain the turn rate through computer control.

My flying experience comes from gliders, which are of course have longer wings and slower roll than a fighter, but in those when you're coming out of a right turn you apply left stick (and coordinate with left rudder) to bring the wings level.

Flying fighters in X Plane, I have to always pull up to get a turn. Is that a sim error? O once read fighters aircraft are not planes, they are rockets with large control surfaces


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Will wing rough surface shift flow separation upstream?
A rough surface takes energy from the boundary layer, so doesn't a slow, "tired" and thick  boundary layer have too little momentum to fight against an adverse pressure gradient?
If the wing surface is very rough or let's say dirty, will this shift flow separation upstream due to the reasons above?
If yes, how can we distinguish "bad" vs "good" turbulent flow, because vortex generators produce turbulent flow to delay separation?

relevant answers can be found here and here.

There is no universally "bad" or "good" turbulent flow. Outside of Hollywood movies or clumsy propaganda, it's always a mixture of good and bad.
Take those vortex generators: They increase the thickness of the boundary layer and, therefore, friction-induced losses, so their added turbulence is mostly bad. But if you have trouble with oscillating shocks or early flow separation on the outer wing, they can remove the last obstacle to certification. For a few minutes in a hours-long flight this added turbulence becomes good and those vortex generators become indispensable.

A rough surface takes energy from the boundary layer, so doesn't a slow, "tired" and thick boundary layer have too little momentum to fight against an adverse pressure gradient?

Yes, it does, but not always. Large, fast airplanes have wing Reynolds numbers where laminar flow is insignificant. Gliders, on the other hand, can be very sensitive to this phenomenon. Modern glider airfoils, on the other hand, take this into consideration and are rather insensitive to added roughness or bugs. Make no mistake, this early transition costs performance, but handling remains largely unchanged with those modern airfoils.

example;if you polish windsurfing fin and make it glossy,you often get problems with  spinout because fin stall,so every fin is matte with sandpaper 400-600. Question is  how much roughness is good and how much is too much?https://www.youtube.com/watch?v=71OEyklxqsk&ab_channel=JeepGarage

@JurgenM That depends entirely on the Reynolds number. The fin is similar to a golf ball and profits from enforced transition, but you don't see dimples on airplane wings.

Sanding submerged parts with 400 sandpaper “catches” water in between the tiny peaks & troughs, and friction during forward motion is then water-on-water instead of plastic on water. So sanding of the fin reduces friction.

@Koyovis That sounds like voodoo physics. Is there any serious proof of this or just surfer hearsay?

@PeterKämpf slalom board bottom:https://lh3.googleusercontent.com/proxy/5GIgzLeepzQ8pCGE208v_RrGfFgCzwCNFK-VF1WUsiH-tvEgbSpYTjz7kwsZeg4bHIAmAKhDWxS735VjIwGzJU2nYP1eJPhzDb2kLxoM5ubVbdTf8eCDgnnQttJkShKF-JYTEfK6Lo0WIurqw7nqUwO3hEJNSJj4x5Z7C4PRkMfn8pT-K648

@PeterKämpf I dont know if Koyovis explantion is correct,but this is true that all race windsurfing boards and fins are matte with sandpaper 400-1000,they are faster than glossy surface,this has been proven last 30 years in slalom/speed racing.Also polished gloss fins has problem with stall(spinout).Because of sanding, bottom of board looks like it was repaired even when buy it new. Does sanding surface works similar as riblets shark skin?

@JurgenM "Does sanding surface works similar as riblets shark skin?" Yes, if sanding direction is in flow direction. Is it?

@PeterKämpf Yes,in flow direction.Before every race riders re-sand board and fin to clear any imperfection of dirty/salt.And race fins has trailing edge sharp as knife!You can cut yourself easy if you are not careful.


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How to perform an operational test on the attitude and heading indicators?
I just need to know how to perfom it. The only thing I know is that you have to remove them and take them to a "lab".

With a smartphone and a suitable app, if that's legal and all you need? I just don' know how accurate and precise those things are

Compass: There will be a compass painted on the ground somewhere at your airfield. Put the aircraft on it and have it point in several directions (every 30° usually). Then make a deviation table. Gyro: take it to the lab and  move it around on some jig that allows to set it at precise angles. See if the indication is correct.


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If we use upper surface blowing and externally blown flap on the same wing will lift correspondingly be very high?
With two engines, one above and one below the wing, will the lift be the sum of two lift increments (produced by upper surface blowing and external blowing)? I am trying to make a jet suit and I am interested in incorporating this phenomenon in it.
When a jet of air blows over or under an airfoil with a deflected flap, it gives an immense amount of lift due to the change in direction of the jet via the Coanda effect. The engine can be placed either above and below and can achieve lift coefficients up to 7 (when any one of the methods - upper surface blowing or externally blown flap - is used)
What if we use two of them at the same time? To which value can the lift coefficient increase?

Is the wing going the same speed? Then lift would be the same, yes?

didn't understand your statement @CrossRoads

Is there really space to have two engines top and bottom of wing?  Review Boeing YC-14 or MD YC-15.

What creates the lift - the wing, or the engines? If the wing, then lift is determined by speed.  If it's the engines, then you have more of a rocket  than a flying suit.

Two Engines Would Probably Be a Waste
First the basics about lift. Lift can be seen as occurring from one of two effects, a pressure difference across surfaces, or in reaction to the acceleration/deflection of the air (or whatever fluid you are moving through). Any case can be calculated from either method and you will come to the same result. You can't exploit a pressure differential without accelerating air, and you can't accelerate air without creating a pressure differential. Since force in these cases is governed by $F=\dot m\Delta v$ and required power is $P = \frac{1}{2} \dot m (v_{out}^2 - v_{in}^2)$. This means that to maximize efficiency you should try to accelerate the largest quantity of air by the smallest amount.
Now let's talk about flow control. Blown flow control systems increase lift in two ways, first is that they help keep the flow attached, and second, especially in propulsion driven systems like you are describing, is by deflecting part of the thrust downwards. Attaching the flow helps to allow greater angles of attack without stalling, once you have have generated this flow attachment, further actuation of flow control will not help generate any more lift by this method, though it may allow further angles of attack. If you are trying to improve low speed characteristics by allowing extreme angles of attack remember that the increase in lift will approximately follow a sin curve as it results from increasing the angle of deflection for flow across the wing, the first little increase in angle will give lots of lift with little drag, the last will give almost no lift for tons of drag if you want to take it as far as 90 degrees. Also lift is related to speed to a power of two, that means you need a lot more deflection as your speed drops.The second method that blown systems generate lift is by deflection, basically you gain some lift by sacrificing thrust. In the end this is no better than just pointing the engines slightly downwards.
If you want to perform flow control you also have to remember how much area on the wing your exhaust is going to effect. Your jet will likely be small in comparison to the total size of the wing.
At the end of the day, it is probably better to make a system that provides some flow control to the entire wing rather than a massive amount of flow control to a small portion of the wing so spread out those engines rather than trying to stack them; and remember, flow control will give no improvement over simple thrust vectoring at a standstill.

I'm not sure this is viable on a wingsuit; it may be too complex or heavy.  However, I do see the concept perfectly viable on a larger type aircraft that may need short take-off/vertical takeoff capability.  Consider just splitting the airflow into two, above and below the wing using a single engine.  Initial engine airflow may have to be different on the wing chord above/below the wings as well.  Study how such a design would compare in efficiency over just pointing the engine airflow downward.

Let's first consider a simpler example. Have two engines on the same side of the wing, close together. Both experience the same effect, so you can still get that lift coefficient of about 7.
That's no surprise; that's how coefficients work. They're multipliers. And by the basic mathematical rule c*(x+y) = c*x + c*y multipliers can be distributed over terms.
Now moving one of the two engines to the other side of the wing doesn't improve the other engine, so the multiplier doesn't improve either.
As for your second question, that's not simple mathematics anymore. Both effects are ultimately due to the deflection of the airflow. As air is forced downwards by the wing, Newton tells us the wing is forced up. That up force is what we call lift. And it's true that you can deflect a single airstream twice. But after the first time, you've already traded some horizontal speed for vertical speed. And it's not the goal to have just vertical speed in the airflow - that's what we call a helicopter.  So the more you deflected the air flow the first time, the less you can deflect further down the streamline.
I can't give you theoretical results here, this will be down to engineering. And of course, a more complex mechanism definitely adds weight and cost.

yeah my doubt is not if the other engine improves it's performance.. I meant to ask, since we are using both of the methods at the same implying twice the amount of flow being directed down, wouldn't lift double(approximately)